Axial compressor

ABSTRACT

An axial compressor includes: a cylindrical casing; a rotor rotatably disposed in the casing; rotor blades arranged on an outer circumferential surface of the rotor around a central axis of the rotor; stationary blades arranged on an inner circumferential surface of the casing at a position adjacent to and behind the rotor blades in an axial direction, each stationary blade having a tip opposing the rotor and having a chord length defined between a leading edge and a trailing edge of the tip; and at least one bleed passage having a bleed opening that opens out in the outer circumferential surface of the rotor. The bleed opening opens at a position located ahead of a position axially spaced rearward from the leading edges of the tips of the stationary blades by one half of the chord length in such a manner that the bleed opening faces toward the tips.

TECHNICAL FIELD

The present invention relates to an axial compressor, and particularlyto an axial compressor equipped with a bleed structure used in gasturbine engines for aircraft or the like.

BACKGROUND ART

Axial compressors used in gas turbine engines for aircraft are designedto operate properly with a large inflow air volume at rated operation(during large-output operation), such as when cruising, which occupies alarge part of the operation time. At non-rated operation, such as whenidling or taxiing, the inflow air volume is small and the inflowcondition differs from the inflow condition at the rated operation.Therefore, at non-rated operation, the vane cascade may not operatestably, rotating stall may occur, the pressure efficiency may decrease,and/or the total pressure loss may become large.

To address such problems, it is known to provide a bleed hole in a partof a fluid passage of a compressor to suppress rotating stall bybleeding part of the compressed air (JP 559-168296A, for example).

However, in the prior art arrangement, the occurrence of rotating stallcannot be avoided sufficiently, and the effect of suppressing thedecrease in the pressure efficiency of the compressor is small.

As a result of earnest research made by the present inventor, it hasbeen found that the effect of suppressing the decrease in the pressureefficiency of the axial compressor by air bleed depends on the positionof the bleed hole with respect to the chord position of the stationaryblade section, and the pressure loss in the axial compressor isfavorably reduced by providing the bleed hole at a particular chordposition (range).

SUMMARY OF THE INVENTION

A primary object of the present invention is to provide an axialcompressor in which a bleed hole is provided at an appropriate positionsuch that the pressure loss in the axial compressor is effectivelyreduced by air bleed.

Means to Accomplish the Task

One embodiment of the present invention provides an axial compressor(36), comprising: a cylindrical casing (14); a rotor (20) rotatablydisposed in the casing; multiple rotor blades (39) arranged on an outercircumferential surface (20B) of the rotor (20) at a predetermined pitcharound a central axis (X) of the rotor; multiple stationary blades (41)arranged on an inner circumferential surface (14A) of the casing (14) ata position adjacent to and behind the rotor blades (39) in an axialdirection of the rotor, each stationary blade having a tip (41A)opposing the rotor and having a chord length (LC) defined between aleading edge (41B) and a trailing edge (41C) of the tip; and at leastone bleed passage (72) having a bleed opening (70) that opens out in theouter circumferential surface (20B) of the rotor (20), wherein the bleedopening (70) opens at a position located ahead of a position axiallyspaced rearward from the leading edges (41B) of the tips (41A) of thestationary blades (41) by one half of the chord length in such a mannerthat the bleed opening faces toward the tips.

According to this arrangement, the air bleed suppresses laminar flowseparation produced in the downstream direction of the stationaryblades, whereby rotating stall becomes less likely to occur and thepressure loss in the axial compressor is effectively reduced.

Preferably, the bleed opening (70) is located axially rearward of theleading edges (41B) of the tips (41A) of the stationary blades (41) by10-20% of the chord length.

According to this arrangement, the pressure loss in the axial compressorair bleed is reduced remarkably by air bleed particularly when theinflow air volume is small.

Preferably, the at least one bleed passage (72) comprises multiple bleedpassages arranged around the central axis (X) of the rotor (20) at aregular pitch.

According to this arrangement, spreading of air vortices produced in thevicinity of the front edges of the stationary blades to the downstreamside of the stationary blades can be prevented evenly all around thecentral axis of the rotor.

Preferably, the at least one bleed passage (72) extends from the bleedopening (70) obliquely rearward at a predetermined angle (θ) relative tothe axial direction of the rotor (20).

According to this arrangement, at non-rated operation, the compressedair containing vortices can flow from the bleed opening to the bleedpassage easily and smoothly, whereby the pressure loss in the axialcompressor is effectively reduced at non-rated operation.

Preferably, the predetermined angle (θ) is in a range from 20 to 40degrees.

According to this arrangement, at non-rated operation, the compressedair containing vortices can flow from the bleed opening to the bleedpassage easily and smoothly, whereby the pressure loss in the axialcompressor is effectively reduced at non-rated operation.

Thus, the axial compressor of the present invention can reduce thepressure loss in the axial compressor effectively by air bleed.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a sectional view showing an overall structure of a gas turbineengine for aircraft including an axial compressor according to oneembodiment of the present invention;

FIG. 2 is a sectional view schematically showing a part of the axialcompressor;

FIG. 3 is a graph showing inflow air volume-pressure loss coefficientcharacteristics; and

FIG. 4 is a graph showing a relationship between total pressure and aspan length.

DESCRIPTION OF THE PREFERRED EMBODIMENT(S)

In the following, an embodiment of an axial compressor according to thepresent invention will be described with reference to FIGS. 1 to 5.

First, an overview of a gas turbine engine (turbofan engine) foraircraft in which the axial compressor of the present embodiment is usedwill be described with reference to FIG. 1.

As shown in FIG. 1, a gas turbine engine 10 includes a substantiallycylindrical outer casing 12 and an inner casing 14 that are arrangedcoaxially. The inner casing 14 rotatably supports a low pressure rotaryshaft (rotor) 20 therein via a front first bearing 16 and a rear firstbearing 18. A tubular high pressure rotary shaft 26 is arranged so as tobe rotatable around an outer circumference of an axially intermediateportion of the low pressure rotary shaft 20. The front portion of thehigh pressure rotary shaft 26 is supported by the inner casing 14 via afront second bearing 22 while the rear portion of the same is supportedby the low pressure rotary shaft 20 via a rear second bearing 24. Thelow pressure rotary shaft 20 and the high pressure rotary shaft 26 arearranged coaxially, and the central axis thereof is denoted by areference sign “X.”

The low pressure rotary shaft 20 includes a substantially conical tipportion 20A that protrudes more forward than the inner casing 14. Anouter circumference of the tip portion 20A is provided with a front fan28 including multiple fan blades 29, which are made of titanium alloy orthe like and arranged to be spaced apart from one another in thecircumferential direction. Multiple stator vanes 30, each having anouter end joined to the outer casing 12 and an inner end joined to theinner casing 14, are arranged on a downstream side of the front fan 28so as to be spaced apart from one another at a predetermined interval inthe circumferential direction. On a downstream side of the stator vanes30, a bypass duct 32 defined between the outer casing 12 and the innercasing 14 to have an annular cross-sectional shape and an aircompression duct (annular fluid passage) 34 defined coaxially (to becoaxial with the central axis X) in the inner casing 14 to have anannular cross-sectional shape are provided in parallel with each other.

An axial compressor 36 is provided in an inlet of the air compressionduct 34. The axial compressor 36 includes two (front and rear) rotorblade tows 38 provided on an outer circumference of the low pressurerotary shaft 20 and two (front and rear) stationary blade rows 40provided in the inner casing 14, such that the rotor blade rows 38 andthe stationary blade rows 40 are arranged adjacent to each other andalternate in the axial direction.

Each of the rotor blade rows 38 includes multiple rotor blades 39 (seeFIG. 2) extending radially outward from an outer circumferential surface20B of the low pressure rotary shaft 20 in a cantilever fashion andarranged around the central axis X of the low pressure rotary shaft 20at a predetermined pitch. Each of the stationary blade rows 40 includesmultiple stationary blades 41 (see FIG. 2) extending radially inwardfrom an inner circumferential surface 14A of the inner casing 14 (seeFIG. 2) in a cantilever fashion and arranged around the central axis Xof the low pressure rotary shaft 20 at a predetermined pitch at aposition adjacent to and behind the corresponding rotor blade row 38 inthe axial direction of the low pressure rotary shaft 20.

A centrifugal compressor 42 is provided in an outlet of the aircompression duct 34. The centrifugal compressor 42 includes impellers 44provided on an outer circumference of the high pressure rotary shaft 26.A stationary blade row 46 is provided in the outlet of the aircompression duct 34 on an upstream side of the impellers 44. Further, adiffuser 50 is provided at an outlet of the centrifugal compressor 42,wherein the diffuser is fixed to the inner casing 14.

On a downstream side of the diffuser 50, a combustion chamber member 54is provided to define a reverse-flow combustion chamber 52 to whichcompressed air is supplied from the diffuser 50. The inner casing 14 isprovided with multiple fuel injection nozzles 56 for injecting fuel intothe reverse-flow combustion chamber 52. The reverse-flow combustionchamber 52 produces high-pressure combustion gas by combusting air-fuelmixture therein. A nozzle guide vane row 58 is provided in an outlet ofthe reverse-flow combustion chamber 52.

On a downstream side of the reverse-flow combustion chamber 52, a highpressure turbine 60 and a low pressure turbine 62 are provided such thatthe combustion gas produced in the reverse-flow combustion chamber 52 isblown thereto. The high pressure turbine 60 includes a high pressureturbine wheel 64 fixed to an outer circumference of the high pressurerotary shaft 26. The low pressure turbine 62 is provided on a downstreamside of the high pressure turbine 60 and includes multiple nozzle guidevane rows 66 fixed to the inner casing 14 and multiple low pressureturbine wheels 68 provided on an outer circumference of the low pressurerotary shaft 20 arranged in an axially alternating manner.

At the start of the gas turbine engine 10, a starter motor (not shown inthe drawings) drives the high pressure rotary shaft 26 to rotate. Oncethe high pressure rotary shaft 26 starts rotating, the air compressed bythe centrifugal compressor 42 is supplied to the reverse-flow combustionchamber 52, and air-fuel mixture combustion takes place in thereverse-flow combustion chamber 52 to produce combustion gas. Thecombustion gas is blown to the high pressure turbine wheel 64 and thelow pressure turbine wheels 68 to rotate the turbine wheels 64, 68.

Thereby, the low pressure rotary shaft 20 and the high pressure rotaryshaft 26 rotate, which causes the front fan 28 to rotate and brings theaxial compressor 36 and the centrifugal compressor 42 into operation,whereby the compressed air is supplied to the reverse-flow combustionchamber 52. Therefore, the gas turbine engine 10 continues to operateafter the starter motor is stopped.

During the operation of the gas turbine engine 10, part of the airsuctioned by the front fan 28 passes through the bypass duct 32 and isblown out rearward, and generates the main thrust particularly in alow-speed flight. The remaining part of the air suctioned by the frontfan 28 is supplied to the reverse-flow combustion chamber 52 and mixedwith the fuel and combusted, and the combustion gas is used to drive thelow pressure rotary shaft 20 and the high pressure rotary shaft 26 torotate before being blown out rearward to generate thrust.

Next, an air bleed structure provided in the axial compressor 36 will bedescribed with reference to FIG. 2.

The low pressure rotary shaft 20 constituting the rotor is formed withmultiple bleed passages 72 arranged around the central axis X of the lowpressure rotary shaft 20 at a regular pitch, wherein each bleed passage72 includes a circular bleed opening 70 that opens out in the outercircumferential surface 20B of the low pressure rotary shaft 20 towardthe air compression duct 34 (compressed air passage).

Part of the compressed air produced by the axial compressor 36 flowsfrom the bleed openings 70 to the bleed passages 72 to be bled to theoutside of the axial compressor 36.

As shown in FIG. 2, each stationary blade 41 has a tip (free end edge)41A opposing the low pressure rotary shaft (rotor) 20, and the tip 41Aof each stationary blade 41 has a leading edge (front end) 41B and atrailing edge (rear end) 41C such that a chord length LC of thestationary blade 41 is defined as a length between the leading edge 41Band the trailing edge 41C.

With respect to the cord position (position in a direction parallel tothe chord of each stationary blade 41 or the axial direction), eachbleed opening 70 opens at a position ahead of a position axially spacedrearward from the leading edges 41B of the tips 41A of the stationaryblades 41 by one half of the chord length LC (0.5LC) in such a mannerthat the bleed opening 70 faces toward the tips 41A of the stationaryblades 41. It is to be noted that the position of the bleed opening 70may be measured as the position of the center of the bleed opening 70.

Preferably, in the embodiment illustrated in FIG. 2, each bleed opening70 is located in a range from a position axially spaced rearward fromthe leading edges 41B of the tips 41A of the stationary blades 41 by 10%of the chord length LC to a position axially spaced rearward from theleading edges 41B by 20% of the chord length LC. Namely, the bleedopening 70 is preferably located in a chord position range of0.1LC-0.2LC with respect to the leading edges 41B.

FIG. 3 is a graph showing a relationship between the inflow air volumeand the pressure loss coefficient for a case where no bleed openings 70are provided and cases where the bleed openings 70 are provided atrespective different chord positions.

In FIG. 3, a characteristic curve A represents the inflow airvolume-pressure loss coefficient characteristics in a case where thebleed openings 70 are located in a range of 0.1LC-0.2LC axially rearwardof the leading edges 41B, a characteristic curve B represents the inflowair volume-pressure loss coefficient characteristics in a case where thebleed openings 70 are located in a range of 0.4LC-0.5LC axially rearwardof the leading edges 41B, a characteristic curve C represents the inflowair volume-pressure loss coefficient characteristics in a case where thebleed openings 70 are located in a range of 0.8LC-0.9LC axially rearwardof the leading edges 41B, and a characteristic curve D represents theinflow air volume-pressure loss coefficient characteristics in a casewhere no bleed openings 70 are provided.

As will be appreciated from this graph, when the bleed openings 70 arelocated ahead of the position axially spaced rearward from the leadingedges 41B by one half of the chord length LC (0.5LC), preferably in arange of 0.1LC-0.2LC axially rearward of the leading edges 41B, thepressure loss coefficient is small, namely, the pressure loss is small.

FIG. 4 is a graph showing a relationship between the span position andthe total pressure for a case where no bleed openings 70 are providedand cases where the bleed openings 70 are provided at respectivedifferent chord positions. In this graph, the position of the outercircumferential surface 20B of the low pressure rotary shaft 20(innermost end position) is represented by a span length of 0, and theposition of the inner circumferential surface 14A of the inner casing 14(outermost end position) is represented by a span length of 1.

As will be appreciated from this graph, provision of the bleed openings70 improves (or suppresses) the decrease in the total pressure or thepressure loss on the side of the span length of 0.

In the present embodiment, because the bleed openings 70 open ahead of aposition axially spaced rearward from the leading edges 41B by 0.5 LC,and are preferably located in the chord position range of 0.1LC-0.2LCaxially rearward of the leading edges 41B, such that the bleed openings70 face toward the tips 41A of the stationary blades 41 from the side ofthe span length of 0, the compressed air produced in the axialcompressor 36 and containing vortices flows from each bleed opening 70to the corresponding bleed passage 72 to be bled efficiently to theoutside of the axial compressor 36, whereby the laminar flow separationthat occurs in the downstream direction of the stationary blades 41 issuppressed. Thereby, in the axial compressor 36 of the presentembodiment, the pressure loss is effectively suppressed so that a highair compression efficiency is achieved.

Since the bleed openings 70 are provided at multiple positions aroundthe central axis X of the low pressure rotary shaft 20 at a regularpitch, air bleed takes place at these positions around the central axisX of the low pressure rotary shaft 20, so that the suppression of thelaminar flow separation produced in the downstream direction of thestationary blades 41 can be achieved evenly all around the central axisX of the low pressure rotary shaft 20. Thereby, the pressure loss in theaxial compressor 36 can be reduced effectively.

Each bleed passage 72 extends from the corresponding bleed opening 70rearward or in the direction of airflow in the axial compressor 36 andobliquely at a predetermined angle θ relative to the outercircumferential surface 20B (or axial direction) of the low pressurerotary shaft 20. Preferably, the angle θ is in a range from 20 to 40degrees.

Because each bleed passage 72 is inclined as described above, thecompressed air can flow from each bleed opening 70 to the correspondingbleed passage 72 easily and smoothly, whereby the pressure loss in theaxial compressor 36 is reduced effectively.

In the foregoing, the present invention has been described in terms ofthe preferred embodiments thereof, but the present invention is notlimited to the foregoing embodiments and various alterations andmodifications may be made as appropriate.

For instance, it is not necessarily indispensable to provide multiplebleed openings 70 (or multiple bleed passages 72), and it may bepossible to provide only a single bleed opening. The bleed openings 70may not be circular, and may be of any other shape such as an ellipse, arectangle, or an oblong circle that extends along the tips of thestationary blades 41.

Also, not all of the structural elements shown in the aboveembodiment(s) are necessarily indispensable and they may be selectivelyused as appropriate without departing from the scope of the presentinvention.

1. An axial compressor, comprising: a cylindrical casing; a rotorrotatably disposed in the casing; multiple rotor blades arranged on anouter circumferential surface of the rotor at a predetermined pitcharound a central axis of the rotor; multiple stationary blades arrangedon an inner circumferential surface of the casing at a position adjacentto and behind the rotor blades in an axial direction of the rotor, eachstationary blade having a tip opposing the rotor and having a chordlength defined between a leading edge and a trailing edge of the tip;and at least one bleed passage having a bleed opening that opens out inthe outer circumferential surface of the rotor, wherein the bleedopening opens at a position located ahead of a position axially spacedrearward from the leading edges of the tips of the stationary blades byone half of the chord length in such a manner that the bleed openingfaces toward the tips.
 2. The axial compressor according to claim 1,wherein the bleed opening is located axially rearward of the leadingedges of the tips of the stationary blades by 10-20% of the chordlength.
 3. The axial compressor according to claim 1, wherein the atleast one bleed passage comprises multiple bleed passages arrangedaround the central axis of the rotor at a regular pitch.
 4. The axialcompressor according to claim 1, wherein the at least one bleed passageextends from the bleed opening obliquely rearward at a predeterminedangle relative to the axial direction of the rotor.
 5. The axialcompressor according to claim 4, wherein the predetermined angle is in arange from 20 to 40 degrees.